Split Resistant Composite Laminate

• US 20140151507A1
• Filed: 12/03/2012
• Published: 06/05/2014
• Est. Priority Date: 12/03/2012
• Status: Active Grant

• Current Assignee: The Boeing Co.
• Original Assignee: The Boeing Co.
• Inventors: Kismarton, Max, Andrews, Francis E., Woodard, Patrick
• Granted Patent: US 9,878,773 B2
• Time in Patent Office: Days
• US Class Current: 244/133
• CPC Class Codes: B32B 5/12 characterised by the relative arrangement of fibres or filaments of different layers, e.g. the fibres or filaments being parallel or perpendicular to each other
• CPC Class Codes: B64C 2001/0072 from composite materials
• CPC Class Codes: B64C 3/26 Construction, shape, or attachment of separate skins, e.g. panels
• CPC Class Codes: Y02T 50/40 Weight reduction
• CPC Class Codes: Y10T 428/24124 Fibers

Abstract

A composite laminate, method of forming same, and use for same are disclosed. One example of a composite laminate has multiple layers or plies (305A-305E) composed of generally parallel reinforcing fibers (315A-315E) embedded in a matrix (305M). The reinforcing fibers have orientations in the ranges of 3 to 8 degrees, −3 to −8 degrees, 10 to 40 degrees, −10 to −40 degrees, and approximately 90 degrees, the orientations being with respect to a predetermined axis (320), such as an axis of tension (T). A method of manufacturing a composite laminate includes laying a resin and fibers having these orientations and then curing the resulting laminate. One example of a use is for the skin on the fuselage or wing of an aircraft.

Claims

1. A composite laminate, comprising:

a matrix; and

reinforcing fibers embedded in the matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

3 to −

8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 10 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

10 to −

40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis.

2. The composite laminate of claim 1 wherein the reinforcing fibers of the first plurality have an orientation of approximately 5 degrees and the reinforcing fibers of the second plurality have an orientation of approximately −

5 degrees.

3. The composite laminate of claim 1 wherein the reinforcing fibers of the third plurality have an orientation of approximately 30 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −

30 degrees.

4. The composite laminate of claim 1 wherein the reinforcing fibers of the third plurality have an orientation of approximately 20 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −

20 degrees.

5. The composite laminate of claim 1 wherein the reinforcing fibers collectively have a total volume, the reinforcing fibers of the first plurality have a first volume, the reinforcing fibers of the second plurality have a second volume, and the first volume and the second volume collectively comprise 30 to 60 percent of the total volume.

6. The composite laminate of claim 1 wherein the reinforcing fibers collectively have a total volume, the reinforcing fibers of the third plurality have a third volume, the reinforcing fibers of the fourth plurality have a fourth volume, and the third volume and the fourth volume collectively comprise 30 to 60 percent of the total volume.

7. The composite laminate of claim 1 wherein the composite laminate is a portion of a skin of an aircraft.

8. The composite laminate of claim 1 wherein the predetermined axis is an axis of tension on the composite laminate.

9. The composite laminate of claim 1 wherein two of the pluralities, both having reinforcing fibers having an identical orientation, are separated by at least one plurality which has reinforcing fibers having a different orientation.

10. The composite laminate of claim 1 wherein the composite laminate has a midplane, and the pluralities are arranged symmetrically about the midplane.

11. A method of manufacturing a composite laminate, comprising:

embedding reinforcing fibers in a matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

3 to −

8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 10 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

10 to −

40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis.

12. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the first plurality to have an orientation of approximately 5 degrees, and positioning the reinforcing fibers of the second plurality to have an orientation of approximately −

5 degrees.

13. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the third plurality to have an orientation of approximately 30 degrees, and positioning the reinforcing fibers of the fourth plurality to have an orientation of approximately −

30 degrees.

14. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the third plurality to have an orientation of approximately 20 degrees, and positioning the reinforcing fibers of the fourth plurality to have orientations of approximately −

20 degrees.

15. The method of claim 11 wherein embedding the reinforcing fibers comprises providing a first volume of the first plurality of reinforcing fibers and providing a second volume of the second plurality of reinforcing fibers, wherein the first volume and the second volume collectively comprise 30 to 60 percent of a total volume of the reinforcing fibers.

16. The method of claim 11 wherein embedding the reinforcing fibers comprises providing a third volume of the third plurality of reinforcing fibers and providing a fourth volume of the fourth plurality of reinforcing fibers, wherein the third volume and the fourth volume collectively comprise 30 to 60 percent of a total volume of the reinforcing fibers.

17. The method of claim 11 and further comprising forming the composite laminate into at least a portion of a skin of a wing of an aircraft.

18. The method of claim 11 wherein the predetermined axis is an axis of tension.

19. The method of claim 11 wherein two of the pluralities of reinforcing fibers, both having reinforcing fibers having an identical orientation, are separated by at least one plurality of reinforcing fibers which has reinforcing fibers having a different orientation.

20. The method of claim 11 wherein the composite laminate has a midplane, and the pluralities of reinforcing fibers are arranged symmetrically about the midplane.

21. An aircraft, comprising:

a fuselage;

a wing assembly operatively connected to the fuselage; and

a composite laminate incorporated into at least a selected portion of the fuselage or the wing assembly, the composite laminate comprising reinforcing fibers embedded in a matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

3 to −

8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 10 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −

10 to −

40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis.

22. The aircraft of claim 21 wherein the wing assembly comprises a wing skin, and wherein the composite laminate is a portion of the wing skin.

23. The aircraft of claim 21 wherein the reinforcing fibers of the first plurality have an orientation of approximately 5 degrees and the reinforcing fibers of the second plurality have an orientation of approximately −

5 degrees.

24. The aircraft of claim 21 wherein the reinforcing fibers of the third plurality have an orientation of approximately 30 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −

30 degrees.

25. The aircraft of claim 21 wherein the reinforcing fibers of the third plurality have an orientation of approximately 20 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −

20 degrees.

26. The aircraft of claim 21 wherein the reinforcing fibers collectively have a total volume, the reinforcing fibers of the first plurality have a first volume, the reinforcing fibers of the second plurality have a second volume, and the first volume and the second volume collectively comprise 30 to 60 percent of the total volume.

27. The aircraft of claim 21 wherein the reinforcing fibers collectively have a total volume, the reinforcing fibers of the third plurality have a third volume, the reinforcing fibers of the fourth plurality have a fourth volume, and the third volume and the fourth volume collectively comprise 30 to 60 percent of the total volume.